U.S. patent application number 10/139934 was filed with the patent office on 2003-03-27 for application of carbon fiber mesh for space and airborne platform applications.
This patent application is currently assigned to Hughes Electronics Corporation. Invention is credited to Lim, Wah L., Thompson, Brian D..
Application Number | 20030057329 10/139934 |
Document ID | / |
Family ID | 24626762 |
Filed Date | 2003-03-27 |
United States Patent
Application |
20030057329 |
Kind Code |
A1 |
Thompson, Brian D. ; et
al. |
March 27, 2003 |
APPLICATION OF CARBON FIBER MESH FOR SPACE AND AIRBORNE PLATFORM
APPLICATIONS
Abstract
A solar array including a carbon fiber mesh substrate is
provided. The solar panels including a series of hybrid structures
formed along the common and continuous substrate by sandwiching the
substrate between a series of discontinuous upper and lower support
layers. In order to construct the solar panels having such hybrid
structure, a series of top support layers or upper face sheets is
disposed on a upper surface of the substrate and between the
folding sections. The solar cells are placed on top of the upper
face sheets. Similarly, a series of lower support layers or lower
face sheets are disposed on a bottom surface of the substrate and
between the folding sections. The folding sections are the regions
where the upper and lower face sheets discontinue and expose the
underlying substrate. The solar array can be folded along the
folding section when a bending force is applied over one of the
hybrid structures.
Inventors: |
Thompson, Brian D.; (Palos
Verdes, CA) ; Lim, Wah L.; (Santa Ana, CA) |
Correspondence
Address: |
HUGHES ELECTRONICS CORPORATION
PATENT DOCKET ADMINISTRATION
BLDG 001 M/S A109
P O BOX 956
EL SEGUNDO
CA
902450956
|
Assignee: |
Hughes Electronics
Corporation
|
Family ID: |
24626762 |
Appl. No.: |
10/139934 |
Filed: |
May 6, 2002 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
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10139934 |
May 6, 2002 |
|
|
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09654924 |
Sep 5, 2000 |
|
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|
6505795 |
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Current U.S.
Class: |
244/172.7 |
Current CPC
Class: |
Y10T 428/249986
20150401; B64G 1/443 20130101; Y10T 428/249928 20150401; Y10T
428/249924 20150401; Y10S 136/292 20130101 |
Class at
Publication: |
244/173 |
International
Class: |
B64G 001/44; F03H
005/00 |
Claims
We claim:
1. A solar array comprising: a mesh substrate, the mesh substrate
being formed from a matrix of resilient fibers; and a plurality of
solar cells disposed on a first surface of the mesh substrate.
2. A solar array comprising: a mesh substrate; said mesh substrate
consisting essentially of a matrix of randomly oriented resilient
fibers and having a first surface; and a plurality of solar cells
disposed on the first surface of the mesh substrate.
3. The solar array of claim 2, wherein said matrix further
comprises a plurality of voids between said fibers.
4. The solar array of claim 2, wherein said mesh substrate is
characterized by a mass of approximately 3 grams per square
meter.
5. A solar array comprising: a mesh substrate; said mesh substrate
consisting essentially of a matrix of carbon fibers; said carbon
fibers being randomly oriented in said matrix; and a plurality of
solar cells disposed on said mesh substrate.
6. The solar array of claim 5, wherein said matrix further
comprises a plurality of voids between said fibers.
7. The solar array of claim 5, wherein said mesh substrate is
characterized by a mass of approximately 3 grams per square
meter.
8. The solar array of claim 5, further comprising a support layer
interposed between said mesh substrate and said plurality of solar
cells.
9. A solar array comprising: a mesh substrate; said mesh substrate
consisting essentially of a matrix of resilient carbon fibers; said
matrix being characterized by a plurality of voids; said carbon
fibers being randomly oriented in said matrix; and a plurality of
solar cells disposed on said mesh substrate.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application claims the benefit of and is a continuation
of U.S. patent application bearing Ser. No. 09/654,924 filed Sep.
5, 2000.
BACKGROUND OF THE INVENTION
[0002] 1. Field of the Invention
[0003] The present invention generally relates to deployable
structures such as solar arrays and, more particularly, to solar
array support systems.
[0004] 2. Description of the Related Art
[0005] In outer space applications, solar arrays are generally
composed of a series of solar panels for generating electrical
power for systems such as a spacecraft or the like. The conversion
of solar energy into electrical energy through solar cells is an
obvious choice for producing power for such systems. The solar
arrays are typically designed in foldable configurations in which
the solar panels, supporting solar cells, are hingeably connected
edge to edge lengthwise by various attachment systems. Such solar
arrays may also include reflectors to concentrate the solar light
upon the solar arrays. Reflectors are attached to the solar panels
widthwise to the opposite ends of each solar panel.
[0006] The solar arrays are generally mounted on deployment yokes
so that they can be extended or retracted from the spacecraft.
During the launch of the spacecraft, the solar arrays are put into
a stowed configuration where the solar arrays are folded in zigzag
fashion against the spacecraft. Once the spacecraft is in outer
space, the solar arrays are deployed into an extended configuration
where the solar panels and the reflectors are folded away into an
operation position in which the solar cells face the sun.
[0007] Such solar arrays must be adequately designed to withstand
the undesirable physical conditions of such space missions so that
they can properly function throughout their life time. Such
undesirable conditions are generally mechanical and thermal
stresses occurring during the launching and during the operation of
the solar arrays. In this respect, the solar panels supporting the
solar cells must be designed to meet the certain thermal and
mechanical stress and strain requirements so as to protect the
solar cells on them. As the solar cells are made of silicon or
gallium arsenide materials, they are brittle. In other words, the
panels function as the mechanical and the thermal support of the
solar arrays.
[0008] Currently, solar panels are constructed from aluminum
honeycomb substrates. In such structures, the honeycomb substrates
are covered with carbon fiber face sheets on upper and lower
surfaces of the honeycomb substrates. The carbon fiber face sheets
stiffen the honeycomb substrate to increase the strength and the
rigidity of the solar panels in stowed or deployed configurations.
However, such aluminum honeycomb base solar panels are heavy in the
context of such space applications. Another drawback involves their
deployment and stowing systems. Such solar arrays require complex
deployment and stowing systems employing tension wires, springs,
hinges and the like to facilitate the deployment and stowing of the
solar arrays.
[0009] As can be seen, there is a need for light weight, high
temperature resistant, stiff and resilient deployable structures
such as solar arrays, reflectors, and thermal blankets.
SUMMARY OF THE INVENTION
[0010] The present invention provides a deployable structure, such
as a solar array system, utilizing a carbon fiber mesh material as
a substrate. Due to its light-weight and flexibility, substrates
containing the carbon fiber mesh material can be applied, as an
example, to deployable solar arrays and reflectors for air borne
vehicles such as satellites for space based applications as well as
to the solar arrays for use in stratospheric platforms of an air
borne vehicle such as an airplane.
[0011] In one aspect of the present invention, a solar array
comprises a mesh substrate and a plurality of solar cells disposed
on a first surface of the mesh substrate. The mesh substrate is
formed from a matrix of resilient fibers.
[0012] These and other features, aspects and advantages of the
present invention will become better understood with reference to
the following drawings, description and claims.
BRIEF DESCRIPTION OF THE DRAWINGS
[0013] FIG. 1 is a perspective view of a satellite using solar
arrays according to an embodiment of the present invention;
[0014] FIGS. 2A-2B are schematic cross-sectional views of the solar
arrays shown in FIG. 1;
[0015] FIG. 3A is a schematic view showing the solar array of FIG.
2A in a fully retracted state;
[0016] FIG. 3B is a schematic view showing the solar array of FIG.
3A in a fully deployed state;
[0017] FIG. 4 is a magnified schematic illustration of the
structure of the carbon fiber mesh substrate of this invention;
[0018] FIG. 5A is a schematic cross sectional view of a solar panel
according to an embodiment of the present invention;
[0019] FIG. 5B is a schematic cross sectional view of a solar panel
according to another embodiment of the present invention;
[0020] FIG. 5C is a schematic cross sectional view of the solar
panel using a carbon fiber mesh composite;
[0021] FIG. 6A is a cross sectional view of an embodiment of the
solar array of the present invention;
[0022] FIGS. 6B-6D are schematic views illustrating deployed and
stowed states of the solar array shown in FIG. 6A;
[0023] FIGS. 7A-7B are schematic views illustrating stratospheric
platforms using an embodiment of the solar array of the present
invention; and
[0024] FIGS. 8A-8C are schematic views illustrating solar
reflectors of the present invention.
DETAILED DESCRIPTION OF THE INVENTION
[0025] Although the present invention is described in the context
of a solar array system utilizing a carbon fiber mesh material as a
substrate, the scope of the invention is not so limited. Due to its
light weight and flexibility, substrates containing the carbon
fiber mesh material can be applied to deployable structures that
support objects. What is meant by "deployable" is a structure that
can be deployed for use and stowed for non-use. Accordingly, these
deployable structures can include thermal blankets, solar arrays
and reflectors for space and land based applications. In
particular, the present invention can be applied to the solar
arrays for use in air borne vehicles such as satellites and
stratospheric platforms of airplanes. In such uses, the carbon
fiber mesh material may be used in combination with other materials
or laminates so as to enhance its applicability and
versatility.
[0026] The carbon fiber mesh substrate of the present invention
provides a material that is lighter and stiffer than the
conventional substrates employing aluminum honeycomb substrates or
cores. Further, the carbon fiber mesh material of the present
invention may withstand very high operation temperatures, as high
as 2500.degree. C. This property significantly alleviates the
localized heating problems in solar arrays. Adding to the
aforementioned advantages, the carbon fiber mesh material of the
present invention is very resilient to bending. If the carbon fiber
mesh material is folded into a compact shape, when released, it
will spring back to its original shape without leaving any fold
line behind. The present invention advantageously employs this
property of the carbon fiber mesh to construct deployable solar
arrays and solar reflectors.
[0027] Reference will now be made to the drawings wherein like
numerals refer to like parts throughout. FIG. 1 exemplifies a
spacecraft 100 comprising a number of solar arrays 102 of the
present invention in a fully deployed configuration. The solar
arrays may be mounted on a side-wall 104 of a body 106 of the
spacecraft 100 using yokes 108. In this embodiment, although it is
possible to use a multiplicity of the solar arrays, for the purpose
of clarity, only two of the solar arrays 102 are shown on the
spacecraft body 106. It is within the scope of this invention that
the solar arrays 102 of the present invention may be used with any
suitable structure including, but not limited to, a stratospheric
platform, satellite, space station or land based applications. The
solar array 102 of the present invention may comprise a series of
solar panels 110 comprising a plurality of solar cells 112 on an
upper side of the solar arrays 102. Solar panels 110 are separated
from one other at a folding section 114. As will be described
below, the folding sections 114 allow panels 110 to be folded
relative to each other in a zigzag fashion when the solar array is
stowed. The solar array 102 of the present embodiment may also
employ solar reflectors to increase the power output of the solar
cells. In this embodiment, a pair of solar reflectors 116 may be
connected to each solar panel 110 using rigid connectors 118 as in
the manner shown in FIG. 1.
[0028] In FIG. 2A, in an exemplary cross-sectional view of the
solar array 102, a number of solar panels 110 are shown separated
by the folding sections 114. The solar panels 110 may be comprised
of a series of hybrid structures formed along a common and
continuous substrate 120 by sandwiching the substrate 120 between a
series of discontinuous upper and lower support layers 122 and 124.
In order to construct the solar panels having such hybrid
structure, a series of upper support layers 122 or upper face
sheets is disposed on a top surface 126 of the substrate 120 and
between the folding sections 114. The solar cells 112 may be placed
on top of the upper face sheets 122. Similarly, a series of lower
support layers 124 or lower face sheets are disposed on a bottom
surface 128 of the substrate 120 and between the folding sections
114. At this point, the folding sections 114 can be further
described as the regions where the upper and lower face sheets 122
and 124 discontinue and expose the underlying substrate 120. As
will be described more fully below, in this embodiment, the
material of the substrate 120 is preferably less stiff than the
hybrid structure of the solar panels 110 which has the substrate
120 interposed between the upper and the lower face sheets 122 and
124.
[0029] Therefore, as illustrated in FIG. 2B, the solar array 102
can be folded along the folding section 114 when a bending force is
applied, for example, in the direction of the arrow 130 and over
one of the stiffened solar panels 110. However, due to the material
properties of the substrate 120, when the bending force is removed,
the solar array springs back to its straight configuration which is
shown in FIG. 2A. In other words, the bending force elastically
deforms the folding section under the deforming force, i.e.,
bending force in this case. Since the deformation of the substrate
material, which is exposed in the folding section 114, is elastic,
i.e., temporary or non-plastic deformation, when the force is
removed the folding section returns to its original shape thereby
deploying the solar array 102 without needing prior art hinge
systems, springs and other deployment equipment. The dimensions of
the substrate 120 depend on the application. Typically, the folding
section 114 is approximately 3 to 4 times the thickness of the
substrate.
[0030] FIGS. 3A and 3B show an implementation of the present
invention. In FIG. 3A, in a stowed configuration, the solar panels
110 of the solar array 102 are folded against the body 106 of the
spacecraft 100 (shown in FIG. 1) in a zigzag fashion by folding
them along the folding sections 114. In the stowed configuration,
the solar panels 110 are folded into a compact structure and
secured (tied down) against the spacecraft until the time of the
deployment. The stiff structure of the solar panels 110 provides a
rigid mounting base for the solar cells 112, which is less likely
to deform and damage the solar cells. Once the spacecraft reaches
the outer space orbit, as shown in FIG. 3B, the solar array 102 of
the present invention is released and the solar panels 110 are
fully extended into a deployed configuration.
[0031] In the preferred embodiment, the substrate 120 comprises
carbon fiber mesh material such as the one available from Energy
Science Laboratory in San Diego, Calif. The carbon fiber mesh
substrate material of the present invention is a substantially
light-weight material, although it is stiff, resistant to
temperature extremes and resilient to bending. As illustrated in
FIG. 4, the carbon fiber mesh material of the present invention is
comprised of a network of randomly oriented carbon fibers 132
linked into a matrix 134. The fiber matrix may also have a
plurality of voids formed between the carbon fibers 132.
Accordingly, the substrate 120 material is distinguishable from the
carbon fiber face sheets described above in the context of the
prior art. Among other things, the substrate 120 material is not in
a woven or unwoven mat form. Further, the substrate 120 material is
neither composed of unidirectional or bidirectional fibers. As a
result of this structure, the carbon fiber mesh material of the
present invention may have a mass of approximately 3 grams per
square meter. Due to its carbon content (carbon is a refractory
material), the carbon fiber mesh material can tolerate temperatures
as high as 2500.degree. C. This property of the carbon fiber mesh
protects the solar arrays from the harmful effects of localized
heating occurring during the operation of the solar arrays.
[0032] As previously mentioned, the carbon fiber mesh material can
be folded under a bending force, but when the force is removed, the
material recovers to its original shape without leaving a fold
line. In this embodiment, the face sheets 122 and 124 may
preferably be carbon fiber mats. Alternatively, Kevlar.TM. mats can
also be used. The carbon fiber mats are generally reinforced carbon
fiber composites impregnated with epoxy. The orientation of the
carbon fibers in the mat is controlled so as to provide maximum
possible tensile strength in a given direction. As will be
described below, multiple layers of face sheets can be used with
different carbon fiber orientations to derive stiffness to bending
along different axes.
[0033] In accordance with the principles of the present invention,
the solar panels 110 of the solar arrays may be fabricated in
various alternative ways. As illustrated in FIG. 5A, in order to
form the above mentioned hybrid panel structure, the upper and
lower carbon fiber face-sheets 122 and 124 are bonded over the top
and bottom surfaces 126 and 128 of the carbon mesh substrate 120
using a bonding material 136. The solar cells may be bonded to the
upper face sheet 122 using the bonding material 136. In this
embodiment a preferred bonding material is epoxy. As previously
mentioned, the face sheets 122 and 124 stiffen the substrate in the
hybrid structure of the solar panels 110.
[0034] FIG. 5B shows another embodiment of forming the hybrid
structure of the solar panels using multiple face sheets. In this
embodiment, after bonding the upper and lower face sheets 122 and
126 to the carbon fiber mesh substrate 120, face sheets 138 and 140
are bonded to the face sheets 122 and 124 respectively. The solar
cells 112 are bonded to the face sheet 138 using the bonding
material 136. As previously mentioned, the use of multiple layers
of face sheets increases the stiffness of the solar panels.
[0035] As shown in FIG. 5C, in another embodiment, a solar panel
142 may comprise a laminated composite structure 144 or carbon
fiber mesh composite comprising a carbon fiber mesh substrate 146
interposed between the carbon fiber mats 148. Solar cells 150 are
bonded to a top side of the carbon fiber mesh composite 144 using a
bonding material 152 such as those used in the previous
embodiments. In this embodiment, it is within the scope of the
present invention that the entire solar array may be formed from
the carbon fiber mesh composite 144 without having individual solar
panels 142.
[0036] As illustrated in FIGS. 6A-6D, for various solar array
applications, the stiffness of the carbon fiber mesh material can
be sufficient without the use of additional face sheet materials.
One such application would be an array composed of amorphous
silicon cells. Since these cells are thinner and more flexible than
conventional solar cells, the carbon fiber mesh alone can provide a
sufficiently stiff substrate. As shown in FIG. 6A, a solar array
160 may comprise a carbon fiber mesh substrate 162 and solar cells
164 bonded on top of the substrate 162 using a bonding material
166. The bonding material 162 is the same bonding material used in
the previous embodiments. FIGS. 6B and 6C show two possible stowing
configurations for the solar array 160, a zigzag and a roll
configuration respectively. As shown in FIG. 6D, when deployed, the
solar array 160 extends into its original shape without having any
fold or fatigue line.
[0037] As shown in FIGS. 7A and 7B, a stratospheric platform 170
may comprise a number of solar arrays 172 to supply solar energy to
an airplane 174. In this embodiment, the stratospheric platform 170
may be comprised of the wings of the airplane 174. As shown in FIG.
7B, the solar array 172 comprising a solar array with the laminated
composite structure 144 described in FIG. 5C, or the solar array
structure 160 described in FIG. 6A can be applied over an upper
skin 178 of the wing 170. The lower mass and the stiffness
properties of the carbon fiber mesh material can allow overall
reduction in the mass of the wing.
[0038] The solar reflectors 116 shown in FIG. 1 can be also
fabricated using carbon fiber mesh material. As shown in FIG. 1,
the solar reflectors are attached to solar panels 110 of the solar
array 102 using the rigid connectors 118. Solar reflectors
concentrate additional light onto the solar cells to increase the
power output of the solar array. In the prior art, the reflectors
are held in tension by a system of springs and rods and attached to
the solar panels using hinge systems. The deployment of such prior
art systems is difficult to implement.
[0039] As illustrated in FIG. 8A, in one embodiment, the reflectors
116 may be comprised of a carbon fiber mesh substrate 180 having a
reflective layer 182 on top of the substrate 180. The reflective
layer may comprise vapor deposited aluminum (VDA) Kapton.TM.
material. As shown in FIG. 8B, in a fully deployed configuration,
the reflectors 116 are attached to the solar panels 110 using rigid
connectors 118, thereby eliminating prior art hinges and tension
wires. As shown in FIG. 8C, the solar reflectors 116 may be stowed
by bending the reflectors toward the top and the bottom of the
solar array 102. The tendency of the carbon fiber mesh to spring
back to its original shape makes it possible to rigidly mount the
reflector to the solar array and stow the reflector by bending the
carbon fiber mesh reflectors.
[0040] It should be understood, of course, that the foregoing
relates to preferred embodiments of the invention and that
modifications may be made without departing from the spirit and
scope of the invention as set forth in the following claims.
* * * * *